This invention relates generally to materials testing and more particularly to a gage to be used for measuring accumulated fatigue damage in a structure which is subjected to repeated loading.
It has long been recognized that due to the repetitious nature of the loads imposed on structures, many parts of the structure suffer fatigue damage, and that the useful life of these parts is limited by the amount of fatigue damage they can withstand and still retain a reasonable margin of safety. It has long been the practice in the industry to identify parts of the structure thought to be fatigue critical in the design phase and to apply various design techniques to minimize fatigue damage. In addition, such parts are often tested in the laboratory under simulated operating conditions to determine their actual fatigue life.
In recent years it has become common to perform fatigue tests on entire assembled airframes by applying thousands of hours of simulated loads, such as aerodynamic loads, and landing gear loads, known to be repetitive in nature to the airframe. Thorough periodic inspections of the test vehicles are made to pinpoint possible fatigue problems. Using these testing techniques, it is normally possible to uncover potential fatigue problem areas and find the solution considerably before any airplane in actual operation would encounter the same problem.
While these testing methods have been generally quite successful in preventing fatigue failures in actual service, they are not helpful in estimating the accumulated fatigue damage or remaining fatigue life of a particular structural component in a particular airplane in the fleet. These methods are based on the assumption that the life history of all airplanes in the fleet can be fairly represented by a statistical approximation of the number of loads of various magnitudes that will be encountered in service, called a fatigue spectrum, and that the application of a cyclic loading pattern based on this fatigue spectrum to a laboratory specimen will result in an amount of fatigue damage equivalent to that which the same part would suffer in actual service when subjected to random loading.
Since different airplanes in the fleet may operate in vastly different climates and may see different types of service, the first assumption is somewhat questionable. As to the second assumption, it is known that variations in the patterns in which repetitive loads are applied to test specimens have a definite effect on the fatigue life of the specimen. Further, it is known that the environment can have a substantial effect on fatigue life and that it may be quite difficult to duplicate environmental conditions in the laboratory to which certain parts are subjected under actual service.
One solution to these problems which has been suggested by others is to place some sort of a gage or indicator directly on a structural member in actual service which will indicate the accumulated fatigue damage suffered by the member and/or the remaining fatigue life after a given period of service.
Among the devices which have been suggested to provide such information is the one described in U.S. Pat. No. 3,272,003 to D. R. Harting dated Sept. 13, 1966. This patent discloses a gage which utilizes a grid of conductive material in the form of a foil or wire which is to be mounted on the structure in question. When the part is subjected to repetitive loading, the electrical resistance of the gage gradually changes, and this change in resistance can be correlated with fatigue damage by performing laboratory tests on the same part.
Another device for accomplishing this purpose which has been suggested in the prior art is a fatigue monitor described in U.S. Pat. No. 3,136,154 to R. H. Christensen dated June 9, 1964. This gage is in the form of a flat elongated strip of material which is fastened at intervals to the specimen in question by some appropriate means such as an adhesive. In between the fastening points the gage is necked down with a notch on either side of the gage in order that the unfastened portions of the gage will be strained during testing to a greater degree than the specimen itself. Stress rising means, usually in the form of holes of various sizes, are located between the notches. When the specimen is subjected to repetitive loading, various unattached sections of the gage will suffer fatigue failures in some sequence depending upon the maximum stress developed in each section. Then, according to the inventor, it is possible to correlate accumulated fatigue damage in a specimen with the number of sections of the gage which have failed at any particular point in a test sequence.
Another technique suggested for monitoring accumulative fatigue damage in a part in actual service is described in U.S. Pat. No. 2,920,480 to T. Haas dated Jan. 12, 1960. In this patent it is suggested that a number of small identical sensing elements be attached to the specimen in question. Prior to attachment however, the elements are "pre-damaged" by subjecting them to a load spectrum which is "equivalent" to the actual service fatigue spectrum for a period of time such that the remaining fatigue life of each sensing element is equal to the safe life assigned to the part. Then it is assumed that when one of the sensing elements attached to the part in actual service fails, the safe fatigue life of the part has been expended.
Other attempts have been made to measure remaining fatigue life by utilizing ultrasonic devices and investigating the physical changes which occur in the material as a result of fatigue. However, most of these previous methods and devices are unsatisfactory for large scale testing of parts in actual service, either because they involve time consuming and elaborate preparation, extensive amounts of wiring, or produce results which are too inaccurate or difficult to interpret. Current practice, therefore, tends toward the use of some device to measure loads or strains experienced by the airframe and then through the use of certain assumptions to predict the relationship between these measured loads and the resulting fatigue damage.